Compact co-axial rotor system for a rotary wing aircraft and a control system therefor

ABSTRACT

A dual, counter rotating, coaxial rotor system provides an upper and lower rotor system, with a reduced axial rotor separation distance along a common axis by way of rotor tip position sensing and rotor position controls to avoid tip contact.

BACKGROUND OF THE INVENTION

The present invention is a divisional application of U.S. patentapplication Ser. No. 10/830,160, filed Apr. 21, 2004.

The present invention relates to a coaxial rotor system, and moreparticularly to a coaxial rotor system with closely spaced articulatedrotors.

Future military forces require enhanced vertical lift capabilities in acompact package. The CH-53E is currently the world's largest shipboardcompatible helicopter. A significant consideration in the design of theCH-53E was shipboard compatibility. The CH-53E effectively defines themaximum aircraft spatial capacity which will fit on the elevators and inthe hangar deck of United States Marine Corps Amphibious Assault Ships,more commonly called an LHA or LHD. Emerging payload weight requirementsare beyond the growth capabilities of the CH-53E while maintainingcurrent shipboard compatibility requirements. Thus, a conventionalhelicopter configuration like the CH-53E would be too large to fit inthe hangar deck or on the elevator of an LHA or LHD.

Conventional coaxial rotor systems are exceeding efficient as liftgenerating mechanism for a heavy lift VTOL aircraft. There are no powerlosses to an anti torque device and rotor efficiency is somewhatimproved relative to a single rotor due to swirl recovery. The aircraftalso has a much lower foot print due to the lack of a tail rotor andsupporting boom structure. Disadvantageously, conventional dual counterrotating coaxial rotor systems require a relatively large separationbetween the rotor systems. This drives the height of a coaxial rotoraircraft to be taller than that of a single rotor aircraft.

Typically the rotor or disks of a conventional dual counter rotatingcoaxial rotor system are axially spaced a distance of approximately 10percent of the rotor diameter. Such a separation is required to provideadequate space for differential rotor blade flapping and bending toassure clearance therebetween regardless of aircraft maneuver. The bladetip position of a conventional coaxial rotor system is determined by thenatural equilibrium of aerodynamic and inertial forces acting on theblade. Since the rotors are counter rotating, many maneuvers causemirror image, or differential, rotor tilt, which reduces tips separationat some point in the rotation. It has been found from decades ofindustry experience that a hub separation distance of 10% of rotordiameter is adequate for most transport types of aircraft.Disadvantageously, application of such rotor spacing to a heavy liftVTOL aircraft which are capable of emerging vertical lift requirementsresult in an aircraft which will likely not meet current shipboardheight compatibility restrictions.

Accordingly, it is desirable to provide an affordable heavy lift VTOLaircraft with low to moderate risk technologies while being compatiblewith current shipboard restrictions.

SUMMARY OF THE INVENTION

The invention described herein utilizes rotor position control through avariety of methods singularly, or in any combination determined to bemost advantageous, to reduce the separation between coaxial rotors andhence reduce the overall height of the aircraft such that the aircraftfits within the hangar deck of an amphibious assault ship.

The dual, counter rotating, coaxial rotor system according to thepresent invention provides an upper and lower rotor system which areseparated by an axial rotor separation distance of approximately 7.5% orless of the rotor system diameter along a common axis. The 7.5 percentseparation distance is a reduction of approximately 25 percent overconventional coaxial rotor systems which are typically separated by atleast 10 percent.

The rotor system requires a blade tip separation clearance between therotor flapping ranges to assure that the rotor blade tips will notcontact. As found from decades of industry experience with conventionalcoaxial rotor systems the minimum tip separation clearance in anymaneuver is approximately 3 percent of the rotor system diameter orapproximately 35 percent of the axial distance between the rotorsystems.

The present invention determines the relative position of the blades ofeach rotor system and independently controls each rotor and or bladesuch that blade flapping and blade elastic bending are significantlyreduced, hence enabling reduced rotor separation.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of this invention will becomeapparent to those skilled in the art from the following detaileddescription of the currently preferred embodiment. The drawings thataccompany the detailed description can be briefly described as follows:

FIG. 1 is a general schematic view of an exemplary coaxial rotary wingaircraft embodiment for use with the present invention;

FIG. 2 is a general schematic view of an exemplary coaxial rotary wingaircraft in a shipboard stored position;

FIG. 3 is graphical representation of a tip separation between rotorsystems of a coaxial rotor system in response to various maneuvers;

FIG. 4 is a block diagram of one embodiment of a rotor control systemaccording to the present invention;

FIG. 5 is a block diagram of a servo-flap embodiment of a rotor controlsystem according to the present invention;

FIG. 6 is a block diagram of the rotor control system of FIG. 4illustrating differential swash plate movement;

FIG. 7 is a block diagram of a swash plateless servo-flap embodiment ofa rotor control system according to the present invention;

FIG. 8 is a block diagram of a tip brake embodiment of a rotor controlsystem according to the present invention;

FIG. 9 is a schematic view of a composite rotor blade embodiment of ablade bending control system according to the present invention;

FIG. 10 is a schematic view of a bend-twist response of a compositerotor blade embodiment of FIG. 9; and

FIG. 11 is a schematic view of a flight control envelope embodiment of arotor control system according to the present invention.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

FIG. 1 schematically illustrates a rotary-wing aircraft 10 having adual, counter rotating, coaxial rotor system 12. The aircraft 10includes an airframe 14 which supports the dual, counter rotating,coaxial rotor system 12 along a common axis A. Although a particularhelicopter configuration is illustrated in the disclosed embodiment,other coaxial propulsor systems which require closely spaced rotors orpropellers in helicopter, airplane and/or tilt rotor type aircraft willalso benefit from the present invention.

The rotor system 12 includes an upper rotor system 16 and a lower rotorsystem 18 which rotate about the common axis A. Each rotor system 16, 18include a multiple of rotor blades 20 mounted to a rotor hub 22, 24.

The rotor systems 16, 18 are separated by an axial rotor separationdistance Sa of less than 10 percent of the rotor system diameter D alongcommon axis A. Preferably, the rotor systems 16, 18 are separated by anaxial distance of approximately 7.5 percent or less of the rotor systemdiameter D along common axis A. The 7.5 percent separation distance is areduction of approximately 25 percent over conventional coaxial rotorsystems which are typically separated by at least 10 percent. Such aseparation distance Sa provides a coaxial configuration which isrelatively compact to fit within a conventional LHA and LHD hanger deckwhile permitting other storage under the folded rotor blades (FIG. 2).

During various maneuvers, the rotor blade tips t of the rotor blades 20will move through a flapping and bending range (defined schematically byarrows f1, f2 in FIG. 1 and by a maneuver chart in FIG. 3). The rotorsystems 16,18 requires a rotor blade tip separation clearance Sc betweenthe flapping range f1, f2 to assure that the rotor blade tips t will notcontact. The separation clearance Sc is preferably 3 percent of therotor system diameter D or approximately 35 percent of the axialdistance Sa between the rotor systems 16,18. It should be understoodthat other clearances will benefit from the present invention; however,the 35 percent separation clearance has proved to be a relativelyconsistent effective separation value as practiced by numerous coaxialrotor system designs.

Referring to FIG. 3, a chart represents rotor tip separation distancesin response to various maneuvers. Applicant has determined that acoaxial rotor system with, for example, a 111 foot diameter rotorsystems 16, 18 separated by only a 7% hub spacing results in rotor tipseparation convergence which may fall below the separation clearance Sc(defined by the dotted line) when utilizing conventional rotor controlsystems. The rotor control systems cited below according to the presentinvention permit a 7% hub spacing while maintaining the 35% separationclearance Sc.

Referring to FIG. 4, each rotor system 16 a, 18 a is independentlycontrolled through a separate swashplate assembly 26 a, 28 a whichselectively articulates each rotor system 16 a, 18 a. Generally, motionof the swash plate assembly 26 a, 28 a along the rotor axis A will causethe rotor blades 20 of the respective rotor system 16 a, 18 a to varypitch collectively and tilting of the swash plate assembly 26 a, 28 awith respect to the axis A will cause the rotor blades 20 to vary pitchcyclically and tilt the rotor disk. The swash plate assemblies 26 a, 28a translate and/or tilt by a separate servo mechanism 30 a, 32 a whichselectively articulates each rotor system 16, 18 independently in bothcyclic and collective in response to a rotor control system 34 a(illustrated schematically). The rotor control system 34 a communicateswith a flight control system 36 which receive pilot inputs from controlssuch as a collective stick, cyclic stick, foot pedals and the like. Asensor suite 38 measures the relative position of the blades 20 on eachrotor system 16, 18 such that the control system 34 determines therelative rotor blade 20 separation. The following paragraphs describethe rotor control methods used to minimize rotor flapping and bendingand are applicable to blades controlled by the actuation methods listedabove.

Referring to FIG. 5, a rotor control system 34 b includes a servo flapcontrolled rotor system 12 b. The servo flap control system is a provenrotor control system in which a rotor blade position is achieved througha rotor blade mounted member. Each blade of rotor systems 16 b, 18 bincludes a partial radius servo flap 42 of a servo flap system 40through, for example only, a series of pushrods and bell cranks that runto relatively small swash plates in the aircraft. It should beunderstood that various actuators and trailing and leading edge slatmounting arrangements will benefit from the present invention. Actuatorssuch as mechanical, electrical, pneumatic, piezoceramic, piezoelectric,hydraulic and the like, both within the blade and external thereto, willalso benefit from the present invention.

Referring to FIG. 6, a rotor control system 34 a independently controlseach swash plate assembly 26 a, 28 a , which selectively controls eachrotor system 16 a, 18 a. That is, the swash plate assembly 26 a, 28 aare not mechanically linked together and are articulated separatelythrough independent servo mechanisms 30 a, 32 a which communicates withthe rotor control system 34 a through a remote communication system suchas a fly-by-wire and/or fly-by-light system. The control system 34 adetermines the relative position of the blades 20 on each rotor system16 a, 18 b and independently controls each swash plate assembly 26 a, 28a to reduce the differential flapping between the rotor systems 16 a, 18a. In other words, if the rotor blade tips of one rotor system 16 a areapproaching the rotor blade tips of the other rotor system 18 a, theswash plates 26 a, 26 b are moved differentially to perform the samemaneuver while increasing separation between the rotor blade tips. Forexample, in forward flight or cruise the rotor blade tips will tendtoward each other on one lateral side of the rotor systems 16 a, 18 a(FIG. 3). By articulating lateral differential cyclic in which swashplate 26 a is moved to a relative positive position from the originalposition while swash plate 26 b is moved to relatively negative positionfrom the original position, the differential swash plate positions areeffectively canceled, forward cyclic is unaffected, and the separationbetween the rotor blade tips is increased (also illustrated by cruisemaneuver point in FIG. 3).

The control system 34 b preferably controls the servo flap system 40 ina higher harmonic control (HHC) methodology. That is, each servo flap 42on each blade 20 of each rotor system 16 b, 18 b (FIG. 7) is pitchedindependently and at a rate greater than once per rotor revolution. Thisenables blade lift corrections as the blade travels around the azimuth.HHC includes the acquiring sensor data in real time and as a result ofthis data, each blade is controlled to provide the desired separationclearance Sc through defined algorithms. In addition the separationclearance Sc may be tailored to multiple performance objectives anddifferent modes of operation, e.g. a high performance mode, a low noisemode, etc., and for different flight conditions and/or configurations,e.g., hover, forward flight, air-to-air engagement, etc. Implementationof HHC may additionally or alternatively include active control and/orprescribed motion functions in response to base conditions, e.g.,forward flight, hover, etc.

Referring to FIG. 8, another rotor control system 34 c controls a tipbrake 46 on each rotor blade 20 c of each rotor system 16 c, 18 c (onlyone illustrated). The tip brakes 46 are preferably split flaps, whichmay be in addition to or integrated with servo flaps 42 (FIG. 7). Thatis, the servo flaps may be split flaps which also operate as tip brakes.Tip brakes are deployed on one rotor and create an unbalanced torquebetween the rotors. The unbalanced torque rotates, or yaws, the aircraftfuselage. The advantage of tip brakes is the zero change in blade lift,and hence no flapping response. Each tip brake 46 is operated inresponse to the control system 34 c to reduce the differential flappingbetween the rotor systems 16 c, 18 c to maintaining the desiredseparation clearance Sc (illustrated by pedal yaw turn maneuver point inFIG. 3).

Referring to FIG. 9, another rotor control system is structurallyintegrated within each rotor blade 20 d by coupling the bend and twistdeflections of the rotor blade through composite design andmanufacturing processes generally understood.

Referring to FIG. 10, preferably, the rotor blades 20 du of the upperrotor system 16 c are designed to increase pitch and hence increaseblade lift when bending downward toward the lower rotor system 18 cwhile the rotor blades 20 dl of the lower rotor system 18 c are designedto decrease pitch and hence reduce lift when bending upward toward theupper rotor system 16 c. The coupling in bending and twisting forces theblades 20 du, 20 dl to maintain the desired separation clearance Sc. Inother words, the blades 20 du, 20 dl will flap away from each other inresponse to a maneuver which causes the blades 20 du, 20 dl to bendtoward each other.

Referring to FIG. 11, another rotor control system 34 e communicateswith the flight control system 36 to selectively restrict portions ofthe rectangular flight envelope E. That is, certain control inputs,which will result in undesirable maneuvers within the flight envelope E,are prevented from occurring so as to maintain the desired separationclearance Sc. FIG. 11 shows typical flight envelope parameters; loadfactor (Nz, measured in g's), airspeed (V, measured in kts), and turnrate (measured in deg/sec). In general maneuvers that may causeexcessive blade bending and flapping are located at the extreme cornersof the envelope. A fly-by-wire (FBW) system enables advanced controllaws where each rotor is controlled independently and will not acceptpilot inputs that could place the coaxial rotor system into a flightstate that may cause the rotor systems to converge and possibly contact.The resulting allowable flight envelope is shown as the truncatedrectangular volume in FIG. 11. Other envelopes including longitudinaland rotational rates and accelerations in each axis can be superimposedto further restrict the aircraft operating envelope to ensure rotor tipclearance Sc. Moreover, FBW and/or HHC application to co-axial rotorsystems permit the reduction or elimination of the tail as well asreduce rotor separation.

Although the rotor control systems disclosed herein are discussedindividually, any rotor control system may be utilized with any other ora multiple of other rotor control systems disclosed herein or otherwiseknown. It should be understood that one rotor control system may bepreferred in one area or maneuver in the flight envelope while anotherrotor control system may be preferred for another area or maneuver inthe flight envelope. Preferably, a combination of rotor control systemsare utilized to assure desired separation clearance Sc throughout theentire flight envelope.

It should be understood that relative positional terms such as“forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like arewith reference to the normal operational attitude of the vehicle andshould not be considered otherwise limiting.

Although particular step sequences are shown, described, and claimed, itshould be understood that steps may be performed in any order, separatedor combined unless otherwise indicated and will still benefit from thepresent invention.

The foregoing description is exemplary rather than defined by thelimitations within. Many modifications and variations of the presentinvention are possible in light of the above teachings. The preferredembodiments of this invention have been disclosed, however, one ofordinary skill in the art would recognize that certain modificationswould come within the scope of this invention. It is, therefore, to beunderstood that within the scope of the appended claims, the inventionmay be practiced otherwise than as specifically described. For thatreason the following claims should be studied to determine the truescope and content of this invention.

1. A coaxial rotor system comprising: a first rotor system having amultiple of first rotor blades which rotates about an axis; a secondrotor system having a multiple of second rotor blades which rotatesabout said axis, said second rotor system spaced an axial distance fromsaid first rotor system; and a rotor control system which independentlycontrols said first rotor system and said second rotor system tomaintain a minimum rotor blade tip separation between said multiple offirst rotor blades and said multiple of second rotor blades.
 2. Thecoaxial rotor system as recited in claim 1, further comprising a rotorblade tip position sensing system in communication with said rotorcontrol system.
 3. The coaxial rotor system as recited in claim 2,further comprising a rotor blade mounted member to independently controleach rotor blade of each of said first rotor system and said secondrotor system in response to said rotor control system and said blade tipposition sensing system.
 4. The coaxial rotor system as recited in claim3, wherein said rotor blade mounted member comprises a servo flap. 5.The coaxial rotor system as recited in claim 3, wherein said rotor blademounted member comprises an aerodynamic tip brake.
 6. The coaxial rotorsystem as recited in claim 1, further comprising a first pitch controlassembly which articulates said first rotor system and a second pitchcontrol assembly which articulates said second rotor system, said firstpitch control assembly controlled independent of said second pitchcontrol assembly in response to said rotor control system.
 7. Thecoaxial rotor system as recited in claim 1, further comprising a firstswash plate which articulates said first rotor system and a second swashplate which articulates said second rotor system, said first swash platecontrolled independent of said second swash plate in response to saidrotor control system.
 8. The coaxial rotor system as recited in claim 1,wherein each of the rotor blades on the first and second rotor systemsincludes a trailing edge servo flap which independently articulates saidfirst rotor system and said second rotor system in response to saidrotor control system.
 9. The coaxial rotor system as recited in claim 1,wherein each of said first and second rotor systems comprise a multipleof bend-twist coupled rotor blades.
 10. The coaxial rotor system asrecited in claim 1, further comprising a flight control system incommunication with said rotor control system to prohibit entry into apredefined portion of a flight envelope which exceeds said minimum rotortip separation.
 11. A coaxial rotor system comprising: a first rotorsystem having at least one servo flap; a second rotor system having atleast one servo flap; said second rotor system being spaced an axialdistance from said first rotor system; and a rotor control system whichindependently controls the at least one servo flap associated with thefirst rotor system and the at least one servo flap associated with thesecond rotor system to maintain a minimum rotor blade tip separationbetween said first rotor system and said second rotor system.
 12. Thecoaxial rotor system as recited in claim 11, wherein at least one ofsaid servo flaps on each of said of said first rotor system and saidsecond rotor system includes a split flap.
 13. The coaxial rotor systemas recited in claim 11, wherein at least one of said servo flaps is anaerodynamic tip brake.
 14. The coaxial rotor system as recited in claim11, wherein at least one of said servo flaps is a trailing edge servoflap.
 15. The coaxial rotor system as recited in claim 11 wherein everyservo flap is a trailing edge servo flap.
 16. The coaxial rotor systemas recited in claim 11, wherein each servo flap on each of said firstand second rotor systems is pitched independently and at a rate greaterthan once per rotor revolution in response to said rotor control systemto maintain said minimum rotor blade tip separation.
 17. The coaxialrotor system as recited in claim 11, wherein said first rotor system andsaid second rotor system are separated by an axial distance ofapproximately 10 percent or less of a diameter of said first and secondrotor system.
 18. A rotary-wing aircraft comprising: a first rotorsystem having a multiple of first rotor blades which rotate about saidaxis, each of said multiple of first rotor blades having at least oneservo flap; a second rotor system having a multiple of second rotorblades which rotate about said axis, said second rotor system spaced anaxial distance from said first rotor system, each of said multiple ofsecond rotor blades having at least one servo flap; a sensor suite whichmeasures a relative position of each of said multiple of first rotorblades and said multiple of second rotor blades; a flight control systemin communication with said sensor suite; and a rotor control systemwhich determines a relative separation distance between each of saidmultiple of first rotor blades and each of said multiple of second rotorblades, said rotor control system in communication with said flightcontrol system to independently control said servo flap on each of saidmultiple of first rotor blades and said multiple of second rotor bladesto maintain a minimum rotor blade tip separation between each of saidmultiple of first rotor blades and each of said multiple of second rotorblades.
 19. The rotary-wing aircraft as recited in claim 18, whereineach servo flap on each of said multiple of first rotor blades and saidmultiple of second rotor blades is pitched independently and at a rategreater than once per rotor revolution in response to said rotor controlsystem to maintain said minimum rotor blade tip separation.
 20. Therotary-wing aircraft as recited in claim 18, wherein said first rotorsystem and said second rotor system are separated by an axial distanceof approximately 10 percent or less of a diameter of said first andsecond rotor system.
 21. The rotary-wing aircraft as recited in claim18, wherein said rotor control system acquires sensor data from saidsensor suite in real time to control said minimum rotor blade tipseparation according to defined algorithms.
 22. The rotary-wing aircraftas recited in claim 21, wherein said defined algorithms are tailored toat least one of a multiple of performance objectives.